The last few pages of this thread (in which I've been lurking for years) have got me thinking. It seems as if the cooling requirements have decreased significantly since the change in engine design from sabre 3 to sabre 4, so much so that frost control is no longer an issue. My question is if this decrease in cooling requirements could allow for a change to methane as a the cryogenic fuel. This could significantly reduce the size and weight of skylon's tanks (and reduce the developmental costs and headaches from dealing with liquid H2). Of course a smaller skylon would make re-entry more difficult - if you believe reaction engines understanding of atmospheric re-entry. I know that the trades around reduced ISP and increased propellant density are complex. However there does seem to be a recent trend, given today's tank building expertise and from lessons learnt from designs like, for example, the x33 to compromise with a lower ISP fuel to gain a benefit from the reduction in weight of tankage and insulation etc.Over a 200 cycle lifetime there would also probably be fairly significant cost savings not only in terms of the cost of fuel but also the systems and infrastructure required to store and pump the fuel at the various launch sites.Is this an obviously answerable question? Or is it something that could go either way and would require a complete vehicle redesign and modelling to answer?In any event I thought it was something worth thinking about - especially given how reaction engines have demonstrated great flexibility in their thought processes by deciding to ditch frost control - a technology which they have spent the better part of twenty years perfecting. It takes a lot of strength of will and character to ditch the fundamental, truly unique part of their business in order to optimise the final vehicle. If they are willing to ditch frost control perhaps they would be willing to ditch liquid hydrogen?
Keep in mind that Skylon's aspect ratio is already below the minimum-drag point, and it can't get any narrower because of the payload bay spec. And since it airbreathes to Mach 5, drag losses are actually fairly substantial - even now I don't think the tankage uses all of the available space (don't quote me on that). I'm not sure a change to a denser propellant would do much, other than reduce the Isp and perhaps complicate the system.I might be wrong - how in-depth were your calculations?
There is the benefit that you've lost a fourth liquid system from the engine which operationally must mean lower maintenance costs.
AFRL is also looking at vehicle concepts.
No, I'm talking about the engine nozzle, the patent for it makes it clear. Pintle or centrebody is what I've read it as in ED papers.Quote"19. A nozzle arrangement according to claim 1, further comprising an actuator arrangement that is arranged to move the second portion of the nozzle between the two positions. "
"19. A nozzle arrangement according to claim 1, further comprising an actuator arrangement that is arranged to move the second portion of the nozzle between the two positions. "
Quote from: lkm on 10/18/2015 06:00 pmNo, I'm talking about the engine nozzle, the patent for it makes it clear. Pintle or centrebody is what I've read it as in ED papers.Quote"19. A nozzle arrangement according to claim 1, further comprising an actuator arrangement that is arranged to move the second portion of the nozzle between the two positions. "Anything is possible but this sounds more like one of those "cover all the bases" clauses that patent drafters like to use. Looking at the drawings you posted you're talking about a movable combustion chamber within an outer combustion chamber. As a way to throttle thrust this seems be a)Mechanically very complex b)Lots of high temperature and pressure engineering combined with lots of cryogenic engineering. c) Large forces would need to be exerted to move one chamber inside another while operating.On the basis that REL likes to avoid unnecessary complexity I think they will stick to throttling by valves on the propellants or hot drive gas (which is inert, rather than super heated steam) supplies. There may be non obvious benefits to moving chamber design, but they'd have to be very substantial to justify the design risk.
FIG. 3 shows a representative one of the nozzles 10 in the rocket mode. In this mode, the second nozzle portion 40 is positioned in the rocket position. In this position, the second nozzle portion 40 is positioned relative to the inner nozzle portion 30 such that the annular throat 50 is closed. In other words, the second nozzle portion 40 is translated to the right in FIG. 3 relative to the first nozzle portion 30. This is such that the generally frusto-conical sections of the two nozzle portions 30, 40 no longer overlap and instead form a contiguous diverging rocket nozzle similar in shape to a conventional rocket nozzle (although it will be noted that the cylindrical section 43 of the second portion 40 still overlaps the first portion 30).
{snip}If you read the patent on SABRE 4 you'll see that in this cycle the precooler exit temperature is designed to never fall below 400K with the bidirectional valve in the cycle directing helium around the precooler to maintain that.
Could nitrogen be used instead of helium at those temperatures?
Quote from: A_M_Swallow on 10/21/2015 01:59 amCould nitrogen be used instead of helium at those temperatures?Maybe, but it would have to reliably not plug a heat exchanger with liquid hydrogen on the other side, even during startup and shutdown. Sounds potentially dicey to me.
a change control board has been established and all of the requirements have been moved from paperspecifications to a computer based requirements tracking programme. .....The formal change control process was also required to enable compliance with the emergingcertification requirements. The board consists of members of the Reaction Engines System engineering team and ESA personnelrepresenting the customers’ interests.
This design is shorter and lighter than previous versions increasing the allowable payload length from 4.4m to 8.6m and the reusable mode GTO payload from 6387kg to 7259kg.
Both these factors were expected to increase the predicted skin temperatures. However the newmodelling also incorporated the results computational fluid dynamics analysis of SKYLON and in practicethe resulting temperatures similar to earlier C1 reentry models.
The multi-foil blanket that lies between the aeroshell and truss structure has been looked at and anew method of construction devised which makes its integration into the vehicle easier and more reliable.Previously it was assumed that each foil layer would be laid one at a time. The new approach creates themonolithic block of foils and spacers all of which are integrated in one piece.
An examination of the SKYLON power and data harness was started in April 2015. This project wastasked to establish a realistic harness mass for the SKYLON airframe. This follows a late realisation thatthe mass of the electrical and optical harness is likely to exceed the total mass of the units that it serves, and therefore it is an important component of system viability.
I understood REL's reluctance to be on making claims based on unproven technologies. SpaceX have demonstrated the use of networking technologies to reduce wiring complexity and weight.
Quote from: JCRM on 10/24/2015 08:16 amI understood REL's reluctance to be on making claims based on unproven technologies. SpaceX have demonstrated the use of networking technologies to reduce wiring complexity and weight.Actually the use of Ethernet and small form factor industrial PC's (104 form factor, IE about 4 inches square) goes back to the Orbital Pegaus. There are are recognized "Avionics Data Networking" standards that are good to at least 100mbs with full dual redundancy and controllable latency and timing variability .The joker in the pack is radiation.ELV's spend most of their time inside the Earths protective magnetic field. They don't complete a single orbit so don't have to face the radiation barrage of the South Atlantic Anomaly and they don't have a planned 200 flight life expectancy, which could cause hard errors to accumulate. The OBC's of the Dragon capsule are about the closest SX have come to this so far. AIUI these rebooted fairly regularly but there were enough of them running to maintain a working capsule at all times. The most demanding systems will likely be the SUS, which will have to perform complex rendezvous motions after exposure to the radiation environment in near GEO orbit.The fact comm sats operate for decades in this environment indicates the problems can be handled, at a price. We will have to see wheather REL goes for a "rad hard by construction" design IE made on a rad hard process, like the 1750A's used by ULA, or a more rad tolerant by design approach using more mainstream parts.
Orion uses Time Triggered Gigabit Ethernet (TTGbE) for all its functions. Here's an article that talks a bit about it (http://mil-embedded.com/articles/orion-avionics-designed-reliability-deep-space/#). The TTGbE technology has now become an industry standard – SAE AS6802 and rad-hard ASICs are available.
What's the main point of using ethernet though? Is it to reduce the number of cables or to increase the data rate?
Lockheed uses wireless sensor networks for DFI on the JSF and has proprietary results on the mass and labor savings from that, along with the benefits for not needing new wall penetrations, etc. We've shared some of that info with Dream Chaser but it could also benefit other programs.
{snip}I think just moving to what is now standard practice in the industry (compared with the SoA when HOTOL was designed) will reduce mass substantially. REL does not appear to like novelty for its own sake. I suspect they will aim to to make the new systems as light as needed but avoid any cutting edge technology where possible. That said since the HOTOL design was last looked at in 1986 things which have a 10 year operating history will still be 20 years more advanced than HOTOL. [EDIT The networking architecture of the recently launched LISA pathfinder mission would seem to be a good model to study as this will be in a radiation environment relatively close to Earth (no dives into the upper Jupiter atmosphere for example) but well outside it's protective magnetic field. Spacewire seems to have been adopted as the hardware standard across all main space agencies ]
Quote from: john smith 19 on 10/26/2015 11:27 pm{snip}I think just moving to what is now standard practice in the industry (compared with the SoA when HOTOL was designed) will reduce mass substantially. REL does not appear to like novelty for its own sake. I suspect they will aim to to make the new systems as light as needed but avoid any cutting edge technology where possible. That said since the HOTOL design was last looked at in 1986 things which have a 10 year operating history will still be 20 years more advanced than HOTOL. [EDIT The networking architecture of the recently launched LISA pathfinder mission would seem to be a good model to study as this will be in a radiation environment relatively close to Earth (no dives into the upper Jupiter atmosphere for example) but well outside it's protective magnetic field. Spacewire seems to have been adopted as the hardware standard across all main space agencies ]In which case REL will be able to buy off the shelf TRL 9 Spacewire components. Obsolescence can be handed by using an interface compatible part; either at the chip or circuit board level.